Cooled cooling air for blade air seal through outer chamber

ABSTRACT

A gas turbine engine according to an example of the present disclosure include a compressor section, a combustor, and a turbine section. The combustor has a radially outer surface that defines a diffuser chamber radially outwardly of the combustor. The turbine section has a high pressure turbine first stage blade that has an outer tip, and a blade outer air seal positioned radially outwardly of the outer tip. A tap for tapping air has been compressed by the compressor and is passed through a heat exchanger. The air downstream of the heat exchanger passes through at least one pipe and into a manifold radially outward of the blade outer air seal, and then passes across the blade outer air seal to cool the blade outer air seal.

BACKGROUND OF THE INVENTION

This application relates to the supply of high pressure cooling air to ablade outer air seal through an outer diameter chamber.

Gas turbine engines are known and typically include a fan delivering airinto a bypass duct as propulsion. The fan also delivers air into acompressor where air is compressed and then delivered into a combustor.The air is mixed with fuel and ignited in the combustor. Products ofthis combustion pass downstream over turbine rotors driving them torotate. The turbine rotors, in turn, rotate compressor rotors and thefan rotor.

As can be appreciated, many components in the turbine section see veryhigh temperatures. Two such components would be the turbine blades andblade outer air seals. Blade outer air seals typically sit radiallyoutwardly of the blades and maintain clearance to increase the efficientuse of the products of combustion.

One type of blade outer air seal is a so-called self-acting clearancecontrol blade outer air seal. In such a blade outer air seal, twocomponents formed of different materials having different coefficientsof thermal expansion combine to control the expansion of the blade outerair seals to, in turn, control the clearance with the blade.

Both the blade and the blade outer air seal are provided with coolingair.

Traditionally, a turbine rotated at the same speed as the fan rotor.More recently, it has been proposed to include a gear reduction betweena fan drive turbine and the fan rotor. With this change, the pressuresand temperatures seen across the turbine sections have increased.

Thus, to drive cooling air into the turbine, the cooling air must be ata higher pressure than in the past. The highest pressure in the gasturbine engine is that downstream of a high pressure compressor.However, this cooling air is also at relatively high temperatures.

Thus, it has been proposed to tap high pressure air from a locationdownstream of the high pressure compressor and pass it through a heatexchanger prior to being delivered to the turbine section for cooling.

It has also been proposed to tap lower pressure from a more forwardstation in the compressor. The air is ducted to a separate auxiliarycompressor and one or more heat exchangers to increase pressure anddecrease the temperature of the air.

SUMMARY OF THE INVENTION

A gas turbine engine according to an example of the present disclosureinclude a compressor section, a combustor, and a turbine section. Thecombustor has a radially outer surface that defines a diffuser chamberradially outwardly of the combustor. The turbine section has a highpressure turbine first stage blade that has an outer tip, and a bladeouter air seal positioned radially outwardly of the outer tip. A tap fortapping air has been compressed by the compressor and is passed througha heat exchanger. The air downstream of the heat exchanger passesthrough at least one pipe and into a manifold radially outward of theblade outer air seal, and then passes across the blade outer air seal tocool the blade outer air seal.

In a further embodiment of any of the foregoing embodiments, the airdownstream of the heat exchanger passes into a mixing chamber and ismixed with higher temperature air from a diffuser chamber outwardly ofthe combustor, and mixed and passes to cool a first stage blade row in ahigh pressure turbine.

In a further embodiment of any of the foregoing embodiments, thecompressor diffuser chamber has an outer boundary defined by an outercore housing and the pipes are radially outward of the outer corehousing.

In a further embodiment of any of the foregoing embodiments, the pipescommunicate with the supply of cooled high pressure air at a locationupstream of the mixing chamber such that air delivered to the manifolddoes not include hot air from the diffuser chamber.

In a further embodiment of any of the foregoing embodiments, themanifold is also outwardly of the outer core housing and communicateswith passages passing through the outer core housing to the blade outerair seal.

In a further embodiment of any of the foregoing embodiments, thepassages passing through the housing extend to a radially outer surfaceof the blade outer air seal and flow in both upstream and downstreamlocations around the blade outer air seal.

In a further embodiment of any of the foregoing embodiments, the airflowing upstream of the blade outer air seal is routed through holes ina seal portion of the blade outer air seal to cool adjacent a leadingedge of the blade outer air seal and the air passing downstream of theblade outer air seal passing through holes in the seal portion of theblade outer air seal to cool adjacent a trailing edge of the blade outerair seal.

In a further embodiment of any of the foregoing embodiments, the bladeouter air seal includes at least two components having different thermalcoefficients of expansion to provide clearance control between an outerperiphery of the blades and an inner periphery of the seal portion.

In a further embodiment of any of the foregoing embodiments, the mixingchamber is radially outward of a compressor diffuser defined downstreamof a downstream most location in a high pressure compressor section andthe air from the mixing chamber passes through vanes in the compressordiffuser.

In a further embodiment of any of the foregoing embodiments, the coolingair is tapped from a location downstream of a downstream most locationin a high pressure compressor.

In a further embodiment of any of the foregoing embodiments, the air istapped from a location upstream of a downstream most location in thecompressor section.

In a further embodiment of any of the foregoing embodiments, themanifold is also outwardly of the outer core housing and communicateswith passages passing through the outer core housing to the blade outerair seal.

In a further embodiment of any of the foregoing embodiments, thecompressor diffuser chamber has an outer boundary defined by an outercore housing and the pipes are radially outward of the outer corehousing.

In a further embodiment of any of the foregoing embodiments, the pipescommunicate with the supply of cooled high pressure air at a locationupstream of the mixing chamber such that air delivered to the manifolddoes not include hot air from the diffuser chamber.

In a further embodiment of any of the foregoing embodiments, themanifold is also outwardly of the outer core housing and communicateswith passages passing through the outer core housing to the blade outerair seal.

In a further embodiment of any of the foregoing embodiments, thepassages passing through the housing extend to a radially outer surfaceof the blade outer air seal and flow in both upstream and downstreamlocations around the blade outer air seal.

In a further embodiment of any of the foregoing embodiments, the airflowing upstream of the blade outer air seal being routed through holesin a seal portion of the blade outer air seal to cool adjacent a leadingedge of the blade outer air seal and the air passing downstream of theblade outer air seal passing through holes in the seal portion of theblade outer air seal to cool adjacent a trailing edge of the blade outerair seal.

In a further embodiment of any of the foregoing embodiments, the mixingchamber is radially outward of a compressor diffuser defined downstreamof a downstream most location in a high pressure compressor section andthe air from the mixing chamber passing through vanes in the compressordiffuser.

In a further embodiment of any of the foregoing embodiments, the bladeouter air seal includes at least two components that have differentthermal coefficients of expansion to provide clearance control betweenan outer periphery of the blades and an inner periphery of the sealportion.

In a further embodiment of any of the foregoing embodiments, a valvecontrols the air passing across the blade outer air seal, and allowscontrol of at least one of an amount, a pressure or a temperature of theair being delivered to the blade outer air seal.

These and other features may be best understood from the followingdrawings and specification.

BRIEF DESCRIPTION OF THE DRAWINGS

FIG. 1 schematically shows a gas turbine engine.

FIG. 2 shows a cooling system.

FIG. 3 shows details of a blade outer air seal cooling system.

DETAILED DESCRIPTION

FIG. 1 schematically illustrates a gas turbine engine 20. The gasturbine engine 20 is disclosed herein as a two-spool turbofan thatgenerally incorporates a fan section 22, a compressor section 24, acombustor section 26 and a turbine section 28. Alternative engines mightinclude an augmentor section (not shown) among other systems orfeatures. The fan section 22 drives air along a bypass flow path B in abypass duct defined within a nacelle 15, while the compressor section 24drives air along a core flow path C for compression and communicationinto the combustor section 26 then expansion through the turbine section28. Although depicted as a two-spool turbofan gas turbine engine in thedisclosed non-limiting embodiment, it should be understood that theconcepts described herein are not limited to use with two-spoolturbofans as the teachings may be applied to other types of turbineengines including three-spool architectures.

The exemplary engine 20 generally includes a low speed spool 30 and ahigh speed spool 32 mounted for rotation about an engine centrallongitudinal axis A relative to an engine static structure 36 viaseveral bearing systems 38. It should be understood that various bearingsystems 38 at various locations may alternatively or additionally beprovided, and the location of bearing systems 38 may be varied asappropriate to the application.

The low speed spool 30 generally includes an inner shaft 40 thatinterconnects a fan 42, a first (or low) pressure compressor 44 and afirst (or low) pressure turbine 46. The inner shaft 40 is connected tothe fan 42 through a speed change mechanism, which in exemplary gasturbine engine 20 is illustrated as a geared architecture 48 to drivethe fan 42 at a lower speed than the low speed spool 30. The high speedspool 32 includes an outer shaft 50 that interconnects a second (orhigh) pressure compressor 52 and a second (or high) pressure turbine 54.A combustor 56 is arranged in exemplary gas turbine 20 between the highpressure compressor 52 and the high pressure turbine 54. A mid-turbineframe 57 of the engine static structure 36 is arranged generally betweenthe high pressure turbine 54 and the low pressure turbine 46. Themid-turbine frame 57 further supports bearing systems 38 in the turbinesection 28. The inner shaft 40 and the outer shaft 50 are concentric androtate via bearing systems 38 about the engine central longitudinal axisA which is collinear with their longitudinal axes.

The core airflow is compressed by the low pressure compressor 44 thenthe high pressure compressor 52, mixed and burned with fuel in thecombustor 56, then expanded through the high pressure turbine 54 and lowpressure turbine 46. The mid-turbine frame 57 includes airfoils 59 whichare in the core airflow path C. The turbines 46, 54 rotationally drivethe respective low speed spool 30 and high speed spool 32 in response tothe expansion. It will be appreciated that each of the positions of thefan section 22, compressor section 24, combustor section 26, turbinesection 28, and fan drive gear system 48 may be varied. For example,gear system 48 may be located aft of combustor section 26 or even aft ofturbine section 28, and fan section 22 may be positioned forward or aftof the location of gear system 48.

The engine 20 in one example is a high-bypass geared aircraft engine. Ina further example, the engine 20 bypass ratio is greater than about six(6), with an example embodiment being greater than about ten (10), thegeared architecture 48 is an epicyclic gear train, such as a planetarygear system or other gear system, with a gear reduction ratio of greaterthan about 2.3 and the low pressure turbine 46 has a pressure ratio thatis greater than about five. In one disclosed embodiment, the engine 20bypass ratio is greater than about ten (10:1), the fan diameter issignificantly larger than that of the low pressure compressor 44, andthe low pressure turbine 46 has a pressure ratio that is greater thanabout five 5:1. Low pressure turbine 46 pressure ratio is pressuremeasured prior to inlet of low pressure turbine 46 as related to thepressure at the outlet of the low pressure turbine 46 prior to anexhaust nozzle. The geared architecture 48 may be an epicycle geartrain, such as a planetary gear system or other gear system, with a gearreduction ratio of greater than about 2.3:1. It should be understood,however, that the above parameters are only exemplary of one embodimentof a geared architecture engine and that the present invention isapplicable to other gas turbine engines including direct driveturbofans.

A significant amount of thrust is provided by the bypass flow B due tothe high bypass ratio. The fan section 22 of the engine 20 is designedfor a particular flight condition—typically cruise at about 0.8 Mach andabout 35,000 feet (10,668 meters). The flight condition of 0.8 Mach and35,000 ft (10,668 meters), with the engine at its best fuelconsumption—also known as “bucket cruise Thrust Specific FuelConsumption (‘TSFC’)”—is the industry standard parameter of lbm of fuelbeing burned divided by lbf of thrust the engine produces at thatminimum point. “Low fan pressure ratio” is the pressure ratio across thefan blade alone, without a Fan Exit Guide Vane (“FEGV”) system. The lowfan pressure ratio as disclosed herein according to one non-limitingembodiment is less than about 1.45. “Low corrected fan tip speed” is theactual fan tip speed in ft/sec divided by an industry standardtemperature correction of [(Tram °R)/(518.7 °R)]^(0.5). The “Lowcorrected fan tip speed” as disclosed herein according to onenon-limiting embodiment is less than about 1150 ft/second (350.5meters/second).

FIG. 2 shows a cooling system 100 for cooling turbine components. Asshown, a compressor section 101 is provided with a tap 102 for tappingpressurized air.

The tap 102 may be at a location upstream from a downstream most portionof a high pressure compressor, in which case, it is typically providedwith a boost compressor to raise its pressure. Alternatively, the aircan be tapped from a location 99 where it has been fully compressed bythe high pressure compressor.

In either case, pressurized air passes through a heat exchanger 104where it is cooled, such as by air. In one embodiment, the heatexchanger 104 may be in the bypass duct as described in FIG. 1. Fromheat exchanger 104, air passes into conduit 106.

From the conduit 106, the air passes into a mixing chamber 108, whichmay be outward of a compressor diffuser 109. The air passes throughvanes in the compressor diffuser 109, such that it is separate from theair downstream of a downstream most compression point 99. The airpasses, as shown at 116, to cool a turbine blade 118. In the mixingchamber 108, hot air is shown at 110 mixing with the cool high pressureair from the conduit 106. This air is from a diffuser chamber 112, andis at the pressure downstream of the downstream most point 99. As such,it mixes easily with the air in the mixing chamber such that the airdelivered at 116 is not unduly cool.

The chamber 112 is outward of a combustion chamber 114. An outer corehousing 113 is positioned outwardly of the chamber 112.

A plurality of pipes 120 (only one of which is shown) tap air from theconduit 106 upstream of the mixing chamber 108. As such, this air isentirely the cooled high pressure cooling air. The air from theplurality of pipes is delivered into a manifold 122 which extendscircumferentially over more than 270° about an axis of rotation of theengine. In embodiments, the manifold 122 extends over 360° about theaxis of rotation. That air then passes through a plurality of pipes 124to cool a blade outer air seal 126.

A valve 150 is shown schematically. The valve 150 may be controlled by acontrol 151 to control the cooling air being sent to the blade outer airseal 126. As an example, the valve may control the amount, pressure ortemperature of the air being delivered to the blade outer air seal 126.An optional line 152 may selectively bypass the heat exchanger 104 toallow temperature control, as an example.

FIG. 3 shows details. As shown, air from the pipes 120 enters manifold122 and then flows through pipe 124 to the blade outer air seal 126. Theblade outer air seal 126 is shown to have components 128 and 130 whichare formed of materials having distinct coefficients of thermalexpansion. These components expand at different rates in response toexposure to heat and provide clearance control for a clearance betweenan inner portion of a seal 132 and an outer tip of the first stage highpressure turbine blade 118 as known.

As shown, the air flows at 134 upstream of the blade outer air sealcomponents 128 and 130 and through holes 137 to cool a leading edge 136of the seal portion 132. This cooling air drives, or controls, theexpansion of the components and thus the clearance control.

Similarly, the air flows at 138 downstream of the blade outer air sealand through holes 142 to cool a trailing edge 140 of the blade outer airseal.

Although an embodiment of this invention has been disclosed, a worker ofordinary skill in this art would recognize that certain modificationswould come within the scope of this invention. For that reason, thefollowing claims should be studied to determine the true scope andcontent of this invention.

1. A gas turbine engine comprising: a compressor section, a combustor,and a turbine section, said combustor having a radially outer surfacedefining a diffuser chamber radially outwardly of said combustor,; saidturbine section including a high pressure turbine first stage bladehaving an outer tip, and a blade outer air seal positioned radiallyoutwardly of said outer tip; a tap for tapping air having beencompressed by said compressor and being passed through a heat exchanger;and said air downstream of said heat exchanger passing through at leastone pipe and into a manifold radially outward of said blade outer airseal, and then passing across said blade outer air seal to cool saidblade outer air seal.
 2. The gas turbine engine as set forth in claim 1,wherein said air downstream of said heat exchanger passes into a mixingchamber and is mixed with higher temperature air from a diffuser chamberoutwardly of said combustor, and mixed and passes to cool a first stageblade row in a high pressure turbine.
 3. The gas turbine engine as setforth in claim 2, wherein said compressor diffuser chamber has an outerboundary defined by an outer core housing and said pipes are radiallyoutward of said outer core housing.
 4. The gas turbine engine as setforth in claim 3, wherein said pipes communicate with the supply ofcooled high pressure air at a location upstream of said mixing chambersuch that air delivered to said manifold does not include hot air fromsaid diffuser chamber.
 5. The gas turbine engine as set forth in claim3, wherein said manifold is also outwardly of said outer core housingand communicates with passages passing through said outer core housingto said blade outer air seal.
 6. The gas turbine engine as set forth inclaim 5, wherein said passages passing through said housing extend to aradially outer surface of said blade outer air seal and flow in bothupstream and downstream locations around said blade outer air seal. 7.The gas turbine engine as set forth in claim 5, wherein said air flowingupstream of said blade outer air seal being routed through holes in aseal portion of said blade outer air seal to cool adjacent a leadingedge of said blade outer air seal and the air passing downstream of saidblade outer air seal passing through holes in said seal portion of saidblade outer air seal to cool adjacent a trailing edge of said bladeouter air seal.
 8. The gas turbine engine as set forth in claim 7,wherein said blade outer air seal includes at least two componentshaving different thermal coefficients of expansion to provide clearancecontrol between an outer periphery of said blades and an inner peripheryof said seal portion.
 9. The gas turbine engine as set forth in claim 7,wherein said mixing chamber is radially outward of a compressor diffuserdefined downstream of a downstream most location in a high pressurecompressor section and said air from said mixing chamber passing throughvanes in said compressor diffuser.
 10. The gas turbine engine as setforth in claim 7, wherein said cooling air is tapped from a locationdownstream of a downstream most location in a high pressure compressor.11. The gas turbine engine as set forth in claim 7, wherein said air istapped from a location upstream of a downstream most location in saidcompressor section.
 12. The gas turbine engine as set forth in claim 2,wherein said manifold is also outwardly of said outer core housing andcommunicates with passages passing through said outer core housing tosaid blade outer air seal.
 13. The gas turbine engine as set forth inclaim 11, wherein said compressor diffuser chamber has an outer boundarydefined by an outer core housing and said pipes are radially outward ofsaid outer core housing.
 14. The gas turbine engine as set forth inclaim 13, wherein said pipes communicate with the supply of cooled highpressure air at a location upstream of said mixing chamber such that airdelivered to said manifold does not include hot air from said diffuserchamber.
 15. The gas turbine engine as set forth in claim 14, whereinsaid manifold is also outwardly of said outer core housing andcommunicates with passages passing through said outer core housing tosaid blade outer air seal.
 16. The gas turbine engine as set forth inclaim 15, wherein said passages passing through said housing extend to aradially outer surface of said blade outer air seal and flow in bothupstream and downstream locations around said blade outer air seal. 17.The gas turbine engine as set forth in claim 16, wherein said airflowing upstream of said blade outer air seal being routed through holesin a seal portion of said blade outer air seal to cool adjacent aleading edge of said blade outer air seal and the air passing downstreamof said blade outer air seal passing through holes in said seal portionof said blade outer air seal to cool adjacent a trailing edge of saidblade outer air seal.
 18. The gas turbine engine as set forth in claim2, wherein said mixing chamber is radially outward of a compressordiffuser defined downstream of a downstream most location in a highpressure compressor section and said air from said mixing chamberpassing through vanes in said compressor diffuser.
 19. The gas turbineengine as set forth in claim 1, wherein said blade outer air sealincludes at least two components having different thermal coefficientsof expansion to provide clearance control between an outer periphery ofsaid blades and an inner periphery of said seal portion.
 20. The gasturbine engine as set forth in claim 19, wherein a valve controls theair passing across said blade outer air seal, and allows control of atleast one of an amount, a pressure or a temperature of the air beingdelivered to the blade outer air seal.